This application relates to a gas turbine engine designed for use on longer range aircraft.
Gas turbine engines are known and may include a fan delivering air into a bypass duct as propulsion air. In addition, the fan typically delivers air into a core housing and to a compressor. There are, typically, at least two compressor rotors with an upstream or lower pressure rotor compressing the air and then delivering it into a downstream or higher pressure rotor. The compressed air from the downstream compressor rotor is typically delivered into a combustion section where it is mixed with fuel and ignited. Products of this combustion may pass downstream over turbine rotors including an upstream turbine rotor that drives the downstream compressor rotor and a downstream turbine rotor that drives the upstream compressor rotor.
In one type engine, the downstream turbine rotor also drove a fan rotor, such that the fan rotor, the upstream compressor rotor, and the downstream turbine rotor all rotated at a single speed. More recently, a gear reduction has been placed between the fan rotor and the downstream turbine rotor or the fan drive turbine.
It is desirable to increase the compression ratio or the amount of compression done to air across the two compressor rotors. However, there has been a significant limitation in that the stress and temperature at the downstream end of the downstream compressor rotor limits how high the overall compression ratio may reach.
This area must be designed to withstand the repeated application of the highest stress situations for the gas turbine engine which typically occurs during take-off.